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United States Patent |
5,233,828
|
Napoli
|
August 10, 1993
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Combustor liner with circumferentially angled film cooling holes
Abstract
A gas turbine engine combustor is provided, having a an annular single wall
sheet metal liner which is generally annular in shape and having disposed
therethrough a multi-hole film cooling means which includes at least one
pattern of small closely spaced film cooling holes angled sharply in the
downstream direction and angled in a circumferential direction wherein the
circumferential angle generally coincides with the swirl angle of the flow
along the surface of the liner. Another embodiment provides a corrugated
aircraft engine sheet metal combustor liner which forms an axially
extending wavy wall to help resist buckling, particularly useful for outer
liners in the combustion section of the engine and in the exhaust section
of gas turbine engines incorporating afterburners.
Inventors:
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Napoli; Phillip D. (West Chester, OH)
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Assignee:
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General Electric Company (Cincinnati, OH)
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Appl. No.:
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951025 |
Filed:
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September 24, 1992 |
Current U.S. Class: |
60/766; 60/755; 60/757 |
Intern'l Class: |
F02C 003/00; F23R 003/06 |
Field of Search: |
60/752,754,755,756,757,267
431/352
|
References Cited
U.S. Patent Documents
2576046 | Nov., 1951 | Scarth | 60/755.
|
3420058 | Jan., 1969 | Howald et al.
| |
3527543 | Sep., 1970 | Howald.
| |
3623711 | Nov., 1971 | Thorstenson.
| |
4232527 | Nov., 1980 | Reider.
| |
4642993 | Feb., 1987 | Sweet.
| |
4653983 | Mar., 1987 | Vehr.
| |
4664597 | May., 1987 | Auxier et al.
| |
4687436 | Aug., 1987 | Shigeta | 431/352.
|
4695247 | Sep., 1987 | Enzaki et al.
| |
4696431 | Sep., 1987 | Buxe.
| |
4773593 | Sep., 1988 | Auxier et al.
| |
4833881 | May., 1989 | Vdoviak et al.
| |
4848081 | Jul., 1989 | Kennedy | 60/261.
|
4878283 | Nov., 1989 | McLean.
| |
4896510 | Jan., 1990 | Foltz.
| |
4923371 | May., 1990 | Ben-Amoz.
| |
Foreign Patent Documents |
90/07087 | Jun., 1990 | WO.
| |
2221979 | Feb., 1990 | GB.
| |
Other References
Multihole Cooling Film Effectiveness and Heat Transfer, by R. E. Mayle and
F. J. Camarata--Transactions of the ASME--Nov., 1975.
Alternate Cooling Configuration for Gas Turbine Combustion Systems, by D.
A. Nealy, S. B. Reider, H. C. Mongia--Allison Gas Turbine Divn., Prepared
by Advisory Group for Aerospace Research & Development 65th Meeting--May
6-10, 1985.
NASA-CR-159656--Advanced Low--Emissions Catalytic--Combustor Program--Phase
I Final Report by G. J. Sturgess--Jun. 1981 report.
|
Primary Examiner: Bertsch; Richard A.
Assistant Examiner: Thorpe; Timothy S.
Attorney, Agent or Firm: Squillaro; Jerome C., Davidson; James P.
Parent Case Text
This application is a continuation of application Ser. No. 07/614,368,
filed Nov. 15, 1990, now abandoned.
Claims
We claim:
1. An annular gas turbine combustor liner for containing a hot combustor
flow, said liner comprising:
a single wall annular shell having a hot surface and a cold surface and at
least one continuous pattern of small closely spaced film cooling holes
angled sharply in the downstream direction from said cold surface to said
hot surface,
said continuous pattern effective to produce a cooling film extending
substantially over the entire length of said shell,
said film cooling holes having a hole diameter, a downstream slant angle,
and spaced at least sufficiently close enough together to effect a cooling
film on said hot surface of said shell during combustor operation, and
said film cooling holes being angled in a circumferential direction.
2. A gas turbine combustor liner as claimed in claim 1 wherein said
circumferential direction coincides with a predetermined swirl angle of
the flow.
3. A gas turbine combustor liner as claimed in claim 1 wherein the
circumferential biased angle is in a range of between 30 and 65 degrees
measured generally from a downstream component of the flow's direction in
the combustor.
4. A gas turbine combustor liner as claimed in claim 1 wherein a portion of
said shell is corrugated to form a shallow wavy wall cross-section.
5. A gas turbine combustor liner as closed in claim 3 wherein said film
cooling holes have a downstream angle slanted from said cold surface of
said shell to said hot surface of said shell and wherein said downstream
angle has a value of about twenty degrees.
6. A gas turbine combustor liner as claimed in claim 3 wherein said film
cooling holes have a downstream angle slanted from said cold surface of
said shell to said hot surface of said shell and wherein said downstream
angle has a value in a range of about twenty degrees.
7. A gas turbine combustor liner as claimed in claim 6 wherein a portion of
said shell is corrugated to form a shallow wavy wall cross-section.
8. An afterburning gas turbine engine exhaust section combustor liner for
containing a hot combustor flow, said exhaust section combustor liner
comprising:
a single wall sheet metal shell having a hot surface and a cold surface
wherein a portion of said shell is corrugated to form a shallow wavy wall
cross-section and
at least one pattern of small closely spaced sharply downstream angled film
cooling holes disposed through said shell having a downstream angle
slanted from said cold surface of said shell to said hot surface of said
shell wherein said downstream angle has a value of about twenty degrees
and said film cooling holes are angled in a circumferential direction,
said continuous pattern effective to produce a cooling film extending
substantially over the entire length of said shell.
9. An afterburning gas turbine engine exhaust section combustor liner as
claimed in claim 8 wherein said circumferential direction coincides with a
predetermined swirl direction of the flow in the combustor.
10. A gas turbine combustor liner as claimed in claim 9 wherein the
circumferential biased angle is in a range of between 30 and 65 degrees
measured generally from a downstream component of the flow's direction in
the combustor.
11. An afterburning gas turbine engine exhaust section combustor liner as
claimed in claim 8 wherein said film cooling holes have a downstream angle
slanted from said cold surface of said shell to said hot surface of said
shell and wherein said downstream angle has a value of about fifteen
degrees.
12. An afterburning gas turbine engine exhaust section combustor liner as
claimed in claim 8 wherein said film cooling holes have a downstream angle
slanted from said cold surface of said shell to said hot surface of said
shell and wherein said downstream angle has a value in a preferred range
of about between ten and twenty degrees.
Description
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates to film cooled combustor liners for use in gas
turbine engines, and more particularly, to aircraft gas turbine engine
combustor liners having cooling holes that are angled in the
circumferential direction.
2. Description of Related Art
Combustor liners are generally used in the combustion section of a gas
turbine engine which is located between the compressor and turbine
sections of the engine. Combustor liners are also used in the exhaust
section of aircraft engines that have afterburners. Combustors generally
include an exterior casing and an interior combustor wherein fuel is
burned producing a hot gas usually at an intensely high temperature such
as 3,000.degree. F. or even higher. To prevent this intense heat from
damaging the combustor before it exits to a turbine, a heat shield or
combustor liner is provided in the interior of the combustor. This
combustor liner thus prevents the intense combustion heat from damaging
the combustor or surrounding engine.
Prior methods for film cooling combustion liners provided circumferentially
disposed rows of film cooling slots such as those depicted in U.S. Pat.
No. 4,566,280 by Burr and U.S..Pat. No. 4,733,538 by Vdoviak et al. which
are typified by complex structures that have non-uniform liner thicknesses
which give rise to thermal gradients which cause low cycle fatigue in the
liner and therefore shorten their potential life expectancy and reduce
their durability. The complex shapes and machining required to produce
these liners negatively effects their cost and weight.
A more detailed discussion of the related art may be found in a related
U.S. patent application Ser. No. 07/614,418 entitled "GAS TURBINE ENGINE
MULTI-HOLE FILM COOLED COMBUSTOR LINER AND METHOD OF MANUFACTURE",
invented by Wakeman et al., filed Nov. 15, 1990, assigned to the same
assignee, and incorporated herein by reference.
Engine designers have long sought to incorporate low weight single wall
combustor liners capable of withstanding the temperatures and pressure
differentials found in combustors. To that end the invention described in
the Wakeman reference provides a single wall, preferably sheet metal,
annular combustor liner having multi-hole film cooling holes which are
disposed through the wall of the liner at sharp downstream angles. The
multi-hole film cooling holes are spaced closely together to form at least
one continuous pattern designed to provide film cooling over the length of
the liner. The present invention provides multi-hole film cooling holes
which have a diameter of about 20 mils with a nominal tolerance of about
.+-.2 mils, are spaced closely together about 61/2 to 71/2 hole diameters
apart, have a downstream angle of 20 degrees with a nominal tolerance of
about .+-.1 degree, and a circumferential angle with respect to the engine
center-line of between 30 and 65 degrees. Axially adjacent holes are
circumferentially offset by half the angle between circumferentially
adjacent holes to further enhance the evenness of the cooling film
injection points. The Wakeman invention further provides an embodiment
wherein the combustor liner may be corrugated so as to form a way wall
which is designed to prevent buckling and is particularly useful for outer
burner liners in the combustion section of gas turbine engines and exhaust
duct burner liners in aircraft gas turbine engines having afterburners.
A phenomena which occurs both in the main combustion section and in the
afterburner combustion section is swirl, wherein swirled patterns of
higher thermal degredation areas are formed on the liner. The patterns
generally coincide with the swirl of the combustor flow induced by
swirlers in the fuel nozzles to promote better combustion and in the
exhaust section by turbine nozzles.
SUMMARY OF THE INVENTION
The present invention provides a multi-hole film cooling means similar to
that described in the Wakeman reference wherein the film cooling
effectiveness is improved by angling or clocking the cooling holes in the
circumferential direction which is best accomplished by drilling the holes
so that the axis of the cooling hole is 30 to 65 degrees to the combustor
flow path. The orientation of clocking is preferrably in the direction of
the combustor swirl pattern as may be generated by the dome swirlers and
stage 1 high pressure turbine nozzle inlets. In accordance with one
embodiment of the present invention, the combustor liner may be corrugated
so as to form a way wall which is designed to prevent buckling and is
particularly useful for outer burner liners in the combustion section of
gas turbine engines and exhaust duct burner liners in aircraft gas turbine
engines having afterburners.
ADVANTAGES
Clocking of the cooling holes in this fashion will impart a tangential
velocity component to the cooling air exiting the multi-hole film cooling
hole and reduce its axial velocity component. This change in velocity
vectors relative to the baseline or unclocked configuration will enhance
the formation of hot side cooling film formed by the hole exit air. This
is accomplished by providing for (1) more uniform film coverage, (2)
increasing film residence time as it traverses the liner gas side flow
path, and (3) reducing the stress concentration factor by aligning the
major axis of the naturally formed elliptical shaped hole exit plane in a
more favorable orientation relative to the engine center-line. This will
also promote better hot gas side film coverage. The present invention
provides a means that allows a reduction in liner cooling flow and
improved performance for pattern and profile limited engines and or
reduced metal temperatures for endurance limited engines.
Combustor liners made in accordance with the present invention dramatically
reduces the radial temperature gradients typically found in conventional
nugget or panel film cooled combustor liners. Reduction of these radial
gradients result in a consequent reduction in thermal hoop stress and
improved Low Cycle Fatigue life. The use of a simple wave form, as found
in conventional augmenting liners, may be used in the outer liner of the
combustion section of the engine, as well as the augmenting liner, to
provide a low cost means of imparting buckling resistance to the combustor
liner.
BRIEF DESCRIPTION OF THE DRAWINGS
The foregoing aspects and other features of the invention are explained in
the following description, taken in connection with the accompanying
drawings where:
FIG. 1 is a diagrammatic view of a typical gas turbine engine including a
core engine combustion section and an afterburning exhaust section having
combustor liners in accordance with the present invention.
FIG. 2 is a perspective view of the core engine combustion section of the
engine depicted in FIG. 1.
FIG. 3 is an enlarged perspective view of a portion of a combustor liner
depicting multi-hole film cooling holes in a portion of a combustor liner
in accordance with the preferred embodiment of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
The gas turbine engine of FIG. 1 represents a conventional aircraft gas
turbine engine having a combustion section combustor and afterburner
employing combustor liners of the present invention.
Referring to FIG. 1, a typical gas turbine engine 10 is shown comprising a
fan section 12 which is in serial flow relationship with an engine core 13
and with a by-pass duct 35 which is generally disposed, in concentric
fashion, about the engine core. Flow from engine core 13 and by-pass duct
35 is discharged to an exhaust section 22 having a nozzle 34 used to help
produce thrust. splitter 17 by-passes a portion of the air flow 27,
referred to as by-pass flow, from fan section 12 through by-pass duct 35
around engine core 13. The remaining airflow, referred to as core air flow
25, is compressed by compressor 14 and discharged to a combustion section
16 which includes axially and circumferentially extending outer and inner
combustor liners 48 and 50, respectively. Outer and inner combustor liners
48 and 50 are radially spaced from each other to define a portion of
annular combustion flow path or combustion zone 33 therebetween where a
portion of core flow 25 is mixed with fuel and the resultant mixture is
combusted. The combustion section produces hot combustion gases which are
mixed with the remainder of the compressor discharge flow and the
resultant heated effluent is then flowed to turbine section 20 which
powers compressor section 14 and fan section 12.
An afterburner 24, as illustrated in FIG. 1, is disposed in exhaust section
22 downstream of turbine section 20 and is operable for burning additional
fuel with bypass air 27 and core flow 25 in order to augment or produce
additional thrust. Thrust augmentation is particularly useful in military
aircraft for short periods of time such as during takeoff, climb and
during combat maneuvers. Exhaust section 22 contains gas flow 32 which is
circumscribed by an annular case 26 and an annular afterburner liner 28
radially inward of casing 26, and a cooling plenum 29 therebetween. The
afterburner may also be referred to as an augmenter. Outer and inner
combustor liners 48 and 50 and afterburner liner 28 provide some generally
similar functions. They contain the hot combustion gases and provide a
flowpath suitable to promote efficient combustion. Pressurized air enters
combustion section 16 where it is mixed with fuel and burned. The hot
gases of combustion, which may in some gas turbine engines exceed
3000.degree. F. exit combustion section 16, flow thereafter past turbine
blades 46 and through the remaining portion of turbine section 20. The hot
gases are then expelled at a high velocity from the engine 10 through
exhaust nozzle 34, whereby the energy remaining therein provides thrust
generation by engine 10.
Referring now to FIG. 2, a perspective view of the combustion section 16 is
depicted comprising a combustor assembly 38 positioned in the compressor
discharge flow 37 between an outer combustor casing 130 and an inner
combustor casing 132 in energized fluid supply communication with the
turbine section denoted by turbine blades 46. Combustor assembly 38 is
further comprised of axially and circumferentially extending outer and
inner combustor liners 48 and 50, respectively, radially spaced from each
other to define a portion of annular flow path or combustion zone 33
therebetween. Outer liner 48 and outer casing 130 form an outer combustor
passage 160 therebetween and inner liner 50 and inner casing 132 form an
inner passage 161 wherein said passages provide for receiving cool
compressor discharge air. Disposed at the upstream end of combustor liners
48 and 50 is a plurality of fuel injectors 52 mounted within a plurality
of apertures 54 in the combustor dome 31 of combustor assembly 38. Note,
that combustor assembly 38 and outer and inner combustor liners 48 and 50
have a preferred annular configuration, extending circumferentially about
the center-line of the engine and dome 31 is of the double dome type to
accommodate the double annular ring of fuel injectors 52. Accordingly,
fuel injectors 52 are circumferentially spaced from each other to provide
a number of injection points for admitting a fuel/air mixture to combustor
assembly 38 over the circumferential extent of annular combustion flow
path 33.
The upstream ends of combustor liners 48 and 50 are formed with means to be
attached to and axially and radially supported by combustor dome 31.
Downstream ends 73 have radial support means such as interference fits or
other conventional support means which provides radial support and allows
for thermal growth of liners 48 and 50.
Outer liner 48 is preferably comprised of a single wall annular sheet or
shell having a generally axially extending generally annular corrugations
60 which provides outer liner 48 with a wavy wall 63 cross-section. Outer
liner 48 has a cold side 57 in contact with the relatively cool air
outside combustion zone 33 and a hot side 61 facing the combustion zone
and includes a means for providing multi-hole film cooling of liner 48.
Referrring to FIG. 3, a frame of reference is provided having axis labelled
X, Y, and Z wherein X indicates the downstream direction of the flow along
the surface of the liner, Y is in the circumferential direction, and Z is
normal to the combustor liner surface on the surface of the liner. The
means for providing multi-hole film cooling, shown in greater detail in
FIG. 3, comprises a plurality of very narrow closely spaced sharply
downstream (indicated by the arrow in FIG. 3) angled film cooling holes 80
which are axially rearward slanted from cold surface 57 to hot surface 61
of liner 48 at an angle in the range of about 15.degree. to 20.degree. and
clocked or slanted in the circumferential direction, indicated by Y in the
frame of reference, at a clock angle B corresponding to the swirl of the
flow which is usually between 30 and 65 degrees with respect to the
downstream direction of the flow indicated by the arrow.
We have found that from a manufacturing and cost standpoint a downstream
slant angle A of about 20.degree. is preferred with respect to either
surface of liner 48. Smaller downstream slant angles A may be may be
advantageous for improved cooling and therefore an alternative downstream
slant angle A in the range of about 20.degree. to 15.degree. may be used
if the associated costs are warranted. Downstream slant angles smaller
than 15 degrees may weaken the liner structure. The holes have a preferred
diameter of 20 mils (0.02 inches) and are preferably spaced about 150 mils
(0.15 inches) off center from each other, as measured between their
respective center-lines 83, or about six and one half (61/2) hole
diameters.
Similarly inner liner 50 is formed of a single wall annular sheet or shell
having a plurality of very narrow closely spaced sharply slanted film
cooling holes 80 which are axially rearward slanted from cold surface 49
to hot surface 51 of liner 50.
Dilution air is primarily introduced by a plurality of circumferentially
extending spaced apart dilution apertures 78 disposed in each of inner and
outer liners 48 and 50. Each aperture 78 and has a cross-sectional area of
substantially greater than the cross-sectional area of one of the
multi-hole cooling holes 80 and are generally far smaller in number.
Dilution apertures 78 and to a smaller extent cooling holes 80 serve to
admit additional air into combustor assembly 38. This additional air mixes
with the air/fuel mixture from injectors 52 and, to some extent, will
promote some additional combustion.
Referring to FIG. 3, liner thickness T, multi-hole film cooling hole
spacing S (the distance between cooling hole center-lines 83), film
cooling hole length L and diameter D, and cooling hole angle A of cooling
holes 80 are a function of the cooling flow requirements to meet the
durability characteristics of the particular engine in which it is used.
Preferably, the combustor liners have a thermal barrier coating on their
hot side 61 to further reduce the heat load into the liners. Cooling holes
80 are laser drilled holes. Typically combustor liner wall thickness T is
sized to meet both mechanical loading requirements and to allow the
cooling flow through cooling hole 80 to develop an adequate length to
diameter ratio (L/D) of least 1.0 and preferably longer. This minimum L/D
is required to form a good film and to maximize convective cooling along
an internal cooling hole surface 81 within cooling hole 80. We have also
found that the cooling holes should be spaced about 7 diameters apart from
each other or between center-lines 83 of adjacent cooling holes 80. The
process of laser drilling is preferably done by drilling the holes from
hot side 61 to cold side 57 of the combustor liner, which for outer liner
48 and afterburner liner 28 is from the inside of the shell out, thereby
producing a diffusion cooling hole having an outlet which is slightly
wider than the cooling hole inlet. The diffusion of the cooling flow
through cooling hole 80 provides a further advantage by enhancing the film
cooling effectiveness which reduces the amount of cooling flow needed
through cooling holes 80 and the pressure and engine performance losses
associated with such cooling means. It may be preferable, particularly in
the case of outer liners 48 resistance means such as corrugations 60 shown
in FIGS. 1 and 2. Buckling of outer 48 liner due to inward pressure load
is a primary design consideration. Small and medium diameter short length
combustors may only require a reasonable liner thickness combined with its
formed shape and end support provided by combustor dome 31 and stator seal
to provide sufficient buckling margin. This margin can be increased by
using significant axial curvature in the liner to increase its section
modulus. Very large combustor liners, having about a 30 inch diameter or
larger, such as outer liner 48 in combustion section 16 and long combustor
liners such as afterburner liner 28 may require additional features to
prevent buckling. The present invention provides corrugations 60 of outer
liner 48 and afterburner liner 28 to restrict the liner deflection and
resist buckling.
The buckling resistance imparted by the wave design of corrugations 60 is
similar to that applied in augmenter liners and must be designed to
provide that the film effectiveness of the liner is not adversely affected
by the wave form. We have found that a shallow wavy form is preferred. An
example of such a shallow wavy wall or corrugated liner is illustrated in
the preferred embodiment which provides, for a combustor section outer
liner 48 having a 30 inch diameter, a trough to crest depth along hot
surface 51 of about 80 mils (0.08 inches) and a crest to crest length of
about 900 mil (0.9 inches) We have found that such a configuration is very
effective for maintaining the integrity of the cooling film and providing
sufficient buckling resistance. The method of manufacturing combustor
liners incorporating the features of the preferred embodiment of the
present invention provides is best described in the above referenced
patent application to Wakeman et al., as applied to an outer liner 48 for
a combustion section 16 having a typical diameter of 30 inches which may
typically contain over 20,000 holes.
While the preferred embodiment of out invention has been described fully in
order to explain its principles, it is understood that various
modifications or alterations may be made to the preferred embodiment
without departing from the scope of the invention as set forth in the
appended claims.
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